[454514566@qq.com]

I want to optimize an airfoil,but I want to set the AOA and airfoil shape as design variables, so I set:
% Optimization design variables, separated by semicolons
DEFINITION_DV= ( 1, 1.0 | airfoil | 0, 0.05 ); ( 1, 1.0 | airfoil | 0, 0.10 ); ( 1, 1.0 | airfoil | 0, 0.15 ); ( 1, 1.0 | airfoil | 0, 0.20 ); ( 1, 1.0 | airfoil | 0, 0.25 ); ( 1, 1.0 | airfoil | 0, 0.30 ); ( 1, 1.0 | airfoil | 0, 0.35 ); ( 1, 1.0 | airfoil | 0, 0.40 ); ( 1, 1.0 | airfoil | 0, 0.45 ); ( 1, 1.0 | airfoil | 0, 0.50 ); ( 1, 1.0 | airfoil | 0, 0.55 ); ( 1, 1.0 | airfoil | 0, 0.60 ); ( 1, 1.0 | airfoil | 0, 0.65 ); ( 1, 1.0 | airfoil | 0, 0.70 ); ( 1, 1.0 | airfoil | 0, 0.75 ); ( 1, 1.0 | airfoil | 0, 0.80 ); ( 1, 1.0 | airfoil | 0, 0.85 ); ( 1, 1.0 | airfoil | 0, 0.90 ); ( 1, 1.0 | airfoil | 0, 0.95 ); ( 1, 1.0 | airfoil | 1, 0.05 ); ( 1, 1.0 | airfoil | 1, 0.10 ); ( 1, 1.0 | airfoil | 1, 0.15 ); ( 1, 1.0 | airfoil | 1, 0.20 ); ( 1, 1.0 | airfoil | 1, 0.25 ); ( 1, 1.0 | airfoil | 1, 0.30 ); ( 1, 1.0 | airfoil | 1, 0.35 ); ( 1, 1.0 | airfoil | 1, 0.40 ); ( 1, 1.0 | airfoil | 1, 0.45 ); ( 1, 1.0 | airfoil | 1, 0.50 ); ( 1, 1.0 | airfoil | 1, 0.55 ); ( 1, 1.0 | airfoil | 1, 0.60 ); ( 1, 1.0 | airfoil | 1, 0.65 ); ( 1, 1.0 | airfoil | 1, 0.70 ); ( 1, 1.0 | airfoil | 1, 0.75 ); ( 1, 1.0 | airfoil | 1, 0.80 ); ( 1, 1.0 | airfoil | 1, 0.85 ); ( 1, 1.0 | airfoil | 1, 0.90 ); ( 1, 1.0 | airfoil | 1, 0.95 ); ( 102, 1.0 | airfoil )

other is the same as the example inv_NACA0012.cfg
and I face a error:
Command = C:\SU2\SU2_CFD config_CFD.cfg
SU2 process returned error '1'
ERROR: Cannot find value AOA in given map.
Please check the name of the variable in the config file.

I hope your rely.thank you

[btracey]

Are you sure you didn't remove the comment from the DV list?

Changing:
% - AOA ( 102, Scale | Markers List )

Into
AOA ( 102, Scale | Markers List )

[454514566@qq.com]

Originally Posted by btracey

Are you sure you didn't remove the comment from the DV list?

Changing:
% - AOA ( 102, Scale | Markers List )

Into
AOA ( 102, Scale | Markers List )

Thank your rely,I want to clear the question again.
The .cfg is showing:


%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
% %
% Stanford University unstructured (SU2) configuration file %
% Case description: Transonic inviscid optimization of a NACA0012 airfoil %
% Author: Francisco Palacios %
% Institution: Stanford University %
% Date: 2013.09.29 %
% File Version 1.0.8 %
% %
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

% ------------- DIRECT, ADJOINT, AND LINEARIZED PROBLEM DEFINITION ------------%
%
% Physical governing equations (EULER, NAVIER_STOKES,
% PLASMA_EULER, PLASMA_NAVIER_STOKES,
% FLUID_STRUCTURE_EULER, FLUID_STRUCTURE_NAVIER_STOKES,
% AEROACOUSTIC_EULER, AEROACOUSTIC_NAVIER_STOKES,
% WAVE_EQUATION, HEAT_EQUATION, LINEAR_ELASTICITY)
PHYSICAL_PROBLEM= EULER
%
% Mathematical problem (DIRECT, ADJOINT, LINEARIZED, ONE_SHOT_ADJOINT)
MATH_PROBLEM= DIRECT
%
% Restart solution (NO, YES)
RESTART_SOL= YES

% -------------------- COMPRESSIBLE FREE-STREAM DEFINITION --------------------%
%
% Mach number (non-dimensional, based on the free-stream values)
MACH_NUMBER= 0.8
%
% Angle of attack (degrees)
AoA= 1.25
%
% Free-stream pressure (101325.0 N/m^2 by default, only Euler flows)
FREESTREAM_PRESSURE= 101325.0
%
% Free-stream temperature (273.15 K by default)
FREESTREAM_TEMPERATURE= 273.15

% -------------- COMPRESSIBLE AND INCOMPRESSIBLE FLUID CONSTANTS --------------%
%
% Ratio of specific heats (1.4 (air), only for compressible flows)
GAMMA_VALUE= 1.4
%
% Specific gas constant (287.87 J/kg*K (air), only for compressible flows)
GAS_CONSTANT= 287.87

% ---------------------- REFERENCE VALUE DEFINITION ---------------------------%
%
% Reference origin for moment computation
REF_ORIGIN_MOMENT_X = 0.25
REF_ORIGIN_MOMENT_Y = 0.00
REF_ORIGIN_MOMENT_Z = 0.00
%
% Reference length for pitching, rolling, and yawing non-dimensional moment
REF_LENGTH_MOMENT= 1.0
%
% Reference area for force coefficients (0 implies automatic calculation)
REF_AREA= 1.0
%
% Reference pressure (101325.0 N/m^2 by default)
REF_PRESSURE= 101325.0
%
% Reference temperature (273.15 K by default)
REF_TEMPERATURE= 273.15
%
% Reference density (1.2886 Kg/m^3 (air), 998.2 Kg/m^3 (water))
REF_DENSITY= 1.2886

% ----------------------- BOUNDARY CONDITION DEFINITION -----------------------%
%
% Marker of the Euler boundary (0 = no marker)
MARKER_EULER= ( airfoil )
%
% Marker of the far field (0 = no marker)
MARKER_FAR= ( farfield )

% ------------------------ SURFACES IDENTIFICATION ----------------------------%
%
% Marker of the surface which is going to be plotted or designed
MARKER_PLOTTING= ( airfoil )
%
% Marker of the surface where the functional (Cd, Cl, etc.) will be evaluated
MARKER_MONITORING= ( airfoil )

% ------------- COMMON PARAMETERS DEFINING THE NUMERICAL METHOD ---------------%
%
% Numerical method for spatial gradients (GREEN_GAUSS, WEIGHTED_LEAST_SQUARES)
NUM_METHOD_GRAD= GREEN_GAUSS
%
% Courant-Friedrichs-Lewy condition of the finest grid
CFL_NUMBER= 10.0
%
% CFL ramp (factor, number of iterations, CFL limit)
CFL_RAMP= ( 1.0, 25, 50.0 )
%
% Runge-Kutta alpha coefficients
RK_ALPHA_COEFF= ( 0.66667, 0.66667, 1.000000 )
%
% Number of total iterations
EXT_ITER= 250

% ------------------------ LINEAR SOLVER DEFINITION ---------------------------%
%
% Linear solver for the implicit (or discrete adjoint) formulation (LU_SGS,
% SYM_GAUSS_SEIDEL, BCGSTAB, GMRES)
LINEAR_SOLVER= FGMRES
%
% Preconditioner of the Krylov linear solver (NONE, JACOBI, LINELET, LUSGS)
LINEAR_SOLVER_PREC= LU_SGS
%
% Min error of the linear solver for the implicit formulation
LINEAR_SOLVER_ERROR= 1E-7
%
% Max number of iterations of the linear solver for the implicit formulation
LINEAR_SOLVER_ITER= 10
%
% Relaxation coefficient
LINEAR_SOLVER_RELAX= 1.0

% -------------------------- MULTIGRID PARAMETERS -----------------------------%
%
% Multi-Grid Levels (0 = no multi-grid)
MGLEVEL= 2
%
% Multi-Grid Cycle (0 = V cycle, 1 = W Cycle)
MGCYCLE= 0
%
% Reduction factor of the CFL coefficient on the coarse levels
MG_CFL_REDUCTION= 0.9
%
% Maximum number of children in the agglomeration stage
MAX_CHILDREN= 250
%
% Maximum length of an agglomerated element (compared with the domain)
MAX_DIMENSION= 0.1
%
% Multi-Grid PreSmoothing Level
MG_PRE_SMOOTH= ( 1, 2, 3, 3 )
%
% Multi-Grid PostSmoothing Level
MG_POST_SMOOTH= ( 0, 0, 0, 0 )
%
% Jacobi implicit smoothing of the correction
MG_CORRECTION_SMOOTH= ( 0, 0, 0, 0 )
%
% Damping factor for the residual restriction
MG_DAMP_RESTRICTION= 1.0
%
% Damping factor for the correction prolongation
MG_DAMP_PROLONGATION= 1.0

% --------------------- FLOW NUMERICAL METHOD DEFINITION ----------------------%
% Convective numerical method (JST, LAX-FRIEDRICH, ROE-1ST_ORDER,
% ROE-2ND_ORDER)
CONV_NUM_METHOD_FLOW= JST
%
% Slope limiter (VENKATAKRISHNAN)
SLOPE_LIMITER_FLOW= VENKATAKRISHNAN
%
% 1st, 2nd and 4th order artificial dissipation coefficients
AD_COEFF_FLOW= ( 0.15, 0.5, 0.02 )
%
% Time discretization (RUNGE-KUTTA_EXPLICIT, EULER_IMPLICIT, EULER_EXPLICIT)
TIME_DISCRE_FLOW= EULER_IMPLICIT

% ---------------- ADJOINT-FLOW NUMERICAL METHOD DEFINITION -------------------%
% Adjoint problem boundary condition (DRAG, LIFT, SIDEFORCE, MOMENT_X,
% MOMENT_Y, MOMENT_Z, EFFICIENCY,
% EQUIVALENT_AREA, NEARFIELD_PRESSURE,
% FORCE_X, FORCE_Y, FORCE_Z, THRUST,
% TORQUE, FREE_SURFACE)
OBJECTIVE_FUNCTION= DRAG
%
% Convective numerical method (JST, LAX-FRIEDRICH, ROE-1ST_ORDER,
% ROE-2ND_ORDER)
CONV_NUM_METHOD_ADJFLOW= JST
%
% Slope limiter (VENKATAKRISHNAN, SHARP_EDGES)
SLOPE_LIMITER_ADJFLOW= VENKATAKRISHNAN
%
% Coefficient for the sharp edges limiter
SHARP_EDGES_COEFF= 3.0
%
% 1st, 2nd, and 4th order artificial dissipation coefficients
AD_COEFF_ADJFLOW= ( 0.15, 0.0, 0.02 )
%
% Time discretization (RUNGE-KUTTA_EXPLICIT, EULER_IMPLICIT)
TIME_DISCRE_ADJFLOW= EULER_IMPLICIT
%
% Reduction factor of the CFL coefficient in the adjoint problem
CFL_REDUCTION_ADJFLOW= 0.8
%
% Limit value for the adjoint variable
LIMIT_ADJFLOW= 1E6
%
% Remove sharp edges from the sensitivity evaluation (NO, YES)
SENS_REMOVE_SHARP= NO
%
% Sensitivity smoothing (NONE, SOBOLEV, BIGRID)
SENS_SMOOTHING= NONE

% --------------------------- PARTITIONING STRATEGY ---------------------------%
%
% Write a paraview file for each partition (NO, YES)
VISUALIZE_PART= NO

% ----------------------- GEOMETRY EVALUATION PARAMETERS ----------------------%
%
% Geometrical evaluation mode (FUNCTION, GRADIENT)
GEO_MODE= FUNCTION
%
% Marker(s) of the surface where geometrical based func. will be evaluated
GEO_MARKER= ( airfoil )

% ----------------------- DESIGN VARIABLE PARAMETERS --------------------------%
%
% Kind of deformation (NO_DEFORMATION, HICKS_HENNE, HICKS_HENNE_NORMAL, PARABOLIC,
% HICKS_HENNE_SHOCK, NACA_4DIGITS, DISPLACEMENT, ROTATION,
% FFD_CONTROL_POINT, FFD_DIHEDRAL_ANGLE, FFD_TWIST_ANGLE,
% FFD_ROTATION)
DV_KIND= HICKS_HENNE
%
% Marker of the surface in which we are going apply the shape deformation
DV_MARKER= ( airfoil )
%
% Parameters of the shape deformation
% - HICKS_HENNE_FAMILY ( Lower(0)/Upper(1) side, x_Loc )
% - NACA_4DIGITS ( 1st digit, 2nd digit, 3rd and 4th digit )
% - PARABOLIC ( 1st digit, 2nd and 3rd digit )
% - DISPLACEMENT ( x_Disp, y_Disp, z_Disp )
% - ROTATION ( x_Orig, y_Orig, z_Orig, x_End, y_End, z_End )
DV_PARAM= ( 1, 0.5 )
%
% Value of the shape deformation deformation
DV_VALUE= 1.0

% ------------------------ GRID DEFORMATION PARAMETERS ------------------------%
%
% Number of smoothing iterations for FEA mesh deformation
DEFORM_LINEAR_ITER= 500
%
% Number of nonlinear deformation iterations (surface deformation increments)
DEFORM_NONLINEAR_ITER= 1
%
% Print the residuals during mesh deformation to the console (YES, NO)
DEFORM_CONSOLE_OUTPUT= YES
%
% Factor to multiply smallest cell volume for deform tolerance (0.001 default)
DEFORM_TOL_FACTOR = 0.001
%
% Type of element stiffness imposed for FEA mesh deformation (INVERSE_VOLUME,
% WALL_DISTANCE, CONSTANT_STIFFNESS)
DEFORM_STIFFNESS_TYPE= INVERSE_VOLUME

% --------------------------- CONVERGENCE PARAMETERS --------------------------%
% Convergence criteria (CAUCHY, RESIDUAL)
%
CONV_CRITERIA= RESIDUAL
%
% Residual reduction (order of magnitude with respect to the initial value)
RESIDUAL_REDUCTION= 6
%
% Min value of the residual (log10 of the residual)
RESIDUAL_MINVAL= -9
%
% Start Cauchy criteria at iteration number
STARTCONV_ITER= 10
%
% Number of elements to apply the criteria
CAUCHY_ELEMS= 100
%
% Epsilon to control the series convergence
CAUCHY_EPS= 1E-6
%
% Direct function to apply the convergence criteria (LIFT, DRAG, NEARFIELD_PRESS)
CAUCHY_FUNC_FLOW= DRAG
%
% Adjoint function to apply the convergence criteria (SENS_GEOMETRY, SENS_MACH)
CAUCHY_FUNC_ADJFLOW= SENS_GEOMETRY
%
% Epsilon for full multigrid method evaluation
FULLMG_CAUCHY_EPS= 1E-3

% ------------------------- INPUT/OUTPUT INFORMATION --------------------------%
% Mesh input file
MESH_FILENAME= mesh_NACA0012_inv.su2
%
% Mesh input file format (SU2, CGNS, NETCDF_ASCII)
MESH_FORMAT= SU2
%
% Mesh output file
MESH_OUT_FILENAME= mesh_out.su2
%
% Restart flow input file
SOLUTION_FLOW_FILENAME= solution_flow.dat
%
% Restart adjoint input file
SOLUTION_ADJ_FILENAME= solution_adj.dat
%
% Output file format (PARAVIEW, TECPLOT)
OUTPUT_FORMAT= TECPLOT
%
% Output file convergence history (w/o extension)
CONV_FILENAME= history
%
% Output file restart flow
RESTART_FLOW_FILENAME= restart_flow.dat
%
% Output file restart adjoint
RESTART_ADJ_FILENAME= restart_adj.dat
%
% Output file flow (w/o extension) variables
VOLUME_FLOW_FILENAME= flow
%
% Output file adjoint (w/o extension) variables
VOLUME_ADJ_FILENAME= adjoint
%
% Output Objective function gradient (using continuous adjoint)
GRAD_OBJFUNC_FILENAME= of_grad.dat
%
% Output file surface flow coefficient (w/o extension)
SURFACE_FLOW_FILENAME= surface_flow
%
% Output file surface adjoint coefficient (w/o extension)
SURFACE_ADJ_FILENAME= surface_adjoint
%
% Writing solution file frequency
WRT_SOL_FREQ= 50
%
% Writing solution file frequency for physical time steps (dual time)
WRT_SOL_FREQ_DUALTIME= 1
%
% Writing convergence history frequency
WRT_CON_FREQ= 1
%
% Writing convergence history frequency (dual time, only written to screen)
WRT_CON_FREQ_DUALTIME= 10
%
% Output rind layers in the solution files
WRT_HALO= NO

% --------------------- OPTIMAL SHAPE DESIGN DEFINITION -----------------------%
% Available flow based objective functions or constraint functions
% DRAG, LIFT, SIDEFORCE, EFFICIENCY,
% FORCE_X, FORCE_Y, FORCE_Z,
% MOMENT_X, MOMENT_Y, MOMENT_Z,
% THRUST, TORQUE, FIGURE_OF_MERIT,
% EQUIVALENT_AREA, NEARFIELD_PRESSURE,
% FREE_SURFACE
%
% Available geometrical based objective functions or constraint functions
% MAX_THICKNESS, 1/4_THICKNESS, 1/2_THICKNESS, 3/4_THICKNESS, AREA, AOA, CHORD,
% MAX_THICKNESS_SEC1, MAX_THICKNESS_SEC2, MAX_THICKNESS_SEC3, MAX_THICKNESS_SEC4, MAX_THICKNESS_SEC5,
% 1/4_THICKNESS_SEC1, 1/4_THICKNESS_SEC2, 1/4_THICKNESS_SEC3, 1/4_THICKNESS_SEC4, 1/4_THICKNESS_SEC5,
% 1/2_THICKNESS_SEC1, 1/2_THICKNESS_SEC2, 1/2_THICKNESS_SEC3, 1/2_THICKNESS_SEC4, 1/2_THICKNESS_SEC5,
% 3/4_THICKNESS_SEC1, 3/4_THICKNESS_SEC2, 3/4_THICKNESS_SEC3, 3/4_THICKNESS_SEC4, 3/4_THICKNESS_SEC5,
% AREA_SEC1, AREA_SEC2, AREA_SEC3, AREA_SEC4, AREA_SEC5,
% AOA_SEC1, AOA_SEC2, AOA_SEC3, AOA_SEC4, AOA_SEC5,
% CHORD_SEC1, CHORD_SEC2, CHORD_SEC3, CHORD_SEC4, CHORD_SEC5
%
% Available design variables
% HICKS_HENNE ( 1, Scale | Mark. List | Lower(0)/Upper(1) side, x_Loc )
% COSINE_BUMP ( 2, Scale | Mark. List | Lower(0)/Upper(1) side, x_Loc, x_Size )
% SPHERICAL ( 3, Scale | Mark. List | ControlPoint_Index, Theta_Disp, R_Disp )
% NACA_4DIGITS ( 4, Scale | Mark. List | 1st digit, 2nd digit, 3rd and 4th digit )
% DISPLACEMENT ( 5, Scale | Mark. List | x_Disp, y_Disp, z_Disp )
% ROTATION ( 6, Scale | Mark. List | x_Axis, y_Axis, z_Axis, x_Turn, y_Turn, z_Turn )
% FFD_CONTROL_POINT ( 7, Scale | Mark. List | Chunk, i_Ind, j_Ind, k_Ind, x_Mov, y_Mov, z_Mov )
% FFD_DIHEDRAL_ANGLE ( 8, Scale | Mark. List | Chunk, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End )
% FFD_TWIST_ANGLE ( 9, Scale | Mark. List | Chunk, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End )
% FFD_ROTATION ( 10, Scale | Mark. List | Chunk, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End )
% FFD_CAMBER ( 11, Scale | Mark. List | Chunk, i_Ind, j_Ind )
% FFD_THICKNESS ( 12, Scale | Mark. List | Chunk, i_Ind, j_Ind )
% FFD_VOLUME ( 13, Scale | Mark. List | Chunk, i_Ind, j_Ind )
% FOURIER ( 14, Scale | Mark. List | Lower(0)/Upper(1) side, index, cos(0)/sin(1) )
%
% Optimization objective function with scaling factor
% ex= Objective * Scale
OPT_OBJECTIVE= DRAG * 0.001
%
% Optimization constraint functions with scaling factors, separated by semicolons
% ex= (Objective = Value ) * Scale, use '>','<','='
OPT_CONSTRAINT= ( LIFT > 0.327 ) * 0.001; ( MAX_THICKNESS > 0.12 ) * 0.001
%
% Optimization design variables, separated by semicolons
DEFINITION_DV= ( 1, 1.0 | airfoil | 0, 0.05 ); ( 1, 1.0 | airfoil | 0, 0.10 ); ( 1, 1.0 | airfoil | 0, 0.15 ); ( 1, 1.0 | airfoil | 0, 0.20 ); ( 1, 1.0 | airfoil | 0, 0.25 ); ( 1, 1.0 | airfoil | 0, 0.30 ); ( 1, 1.0 | airfoil | 0, 0.35 ); ( 1, 1.0 | airfoil | 0, 0.40 ); ( 1, 1.0 | airfoil | 0, 0.45 ); ( 1, 1.0 | airfoil | 0, 0.50 ); ( 1, 1.0 | airfoil | 0, 0.55 ); ( 1, 1.0 | airfoil | 0, 0.60 ); ( 1, 1.0 | airfoil | 0, 0.65 ); ( 1, 1.0 | airfoil | 0, 0.70 ); ( 1, 1.0 | airfoil | 0, 0.75 ); ( 1, 1.0 | airfoil | 0, 0.80 ); ( 1, 1.0 | airfoil | 0, 0.85 ); ( 1, 1.0 | airfoil | 0, 0.90 ); ( 1, 1.0 | airfoil | 0, 0.95 ); ( 1, 1.0 | airfoil | 1, 0.05 ); ( 1, 1.0 | airfoil | 1, 0.10 ); ( 1, 1.0 | airfoil | 1, 0.15 ); ( 1, 1.0 | airfoil | 1, 0.20 ); ( 1, 1.0 | airfoil | 1, 0.25 ); ( 1, 1.0 | airfoil | 1, 0.30 ); ( 1, 1.0 | airfoil | 1, 0.35 ); ( 1, 1.0 | airfoil | 1, 0.40 ); ( 1, 1.0 | airfoil | 1, 0.45 ); ( 1, 1.0 | airfoil | 1, 0.50 ); ( 1, 1.0 | airfoil | 1, 0.55 ); ( 1, 1.0 | airfoil | 1, 0.60 ); ( 1, 1.0 | airfoil | 1, 0.65 ); ( 1, 1.0 | airfoil | 1, 0.70 ); ( 1, 1.0 | airfoil | 1, 0.75 ); ( 1, 1.0 | airfoil | 1, 0.80 ); ( 1, 1.0 | airfoil | 1, 0.85 ); ( 1, 1.0 | airfoil | 1, 0.90 ); ( 1, 1.0 | airfoil | 1, 0.95 );( 102, 1.0 | airfoil )


the erroris showing:
D:\>shape_optimization.py -f inv_NACA0012.cfg -p 0
found: mesh_NACA0012_inv.su2
New Project: ./
Warning, removing old designs... now
New Design: DESIGNS/DSN_001
./DESIGNS/DSN_001
./DESIGNS/DSN_001
Evaluate Inequality Constraints
Lift... the command: C:\SU2\SU2_CFD config_CFD.cfg
the location: D:\DESIGNS\DSN_001\DIRECT
Traceback (most recent call last):
File "C:\SU2\shape_optimization.py", line 124, in <module>
main()
File "C:\SU2\shape_optimization.py", line 69, in main
options.step )
File "C:\SU2\shape_optimization.py", line 107, in shape_optimization
SU2.opt.SLSQP(project,x0,xb,its)
File "C:\SU2\SU2\opt\scipy_tools.py", line 102, in scipy_slsqp
epsilon = 1.0e-06 )
File "C:\Python27\lib\site-packages\scipy\optimize\slsqp.py", line 206, in fm
n_slsqp
constraints=cons, **opts)
File "C:\Python27\lib\site-packages\scipy\optimize\slsqp.py", line 308, in _m
nimize_slsqp
mieq = sum(map(len, [atleast_1d(c['fun'](x, *c['args'])) for c in cons['ine
']]))
File "C:\SU2\SU2\opt\scipy_tools.py", line 187, in con_cieq
cons = project.con_cieq(x)
File "C:\SU2\SU2\opt\project.py", line 223, in con_cieq
return self._eval(konfig, func,dvs)
File "C:\SU2\SU2\opt\project.py", line 172, in _eval
vals = design._eval(func,*args)
File "C:\SU2\SU2\eval\design.py", line 132, in _eval
vals = eval_func(*inputs)
File "C:\SU2\SU2\eval\design.py", line 422, in con_cieq
func = su2func(this_con,config,state)
File "C:\SU2\SU2\eval\functions.py", line 75, in function
aerodynamics( config, state )
File "C:\SU2\SU2\eval\functions.py", line 215, in aerodynamics
info = su2run.direct(config)
File "C:\SU2\SU2\run\direct.py", line 75, in direct
SU2_CFD(konfig)
File "C:\SU2\SU2\run\interface.py", line 93, in CFD
run_command( the_Command )
File "C:\SU2\SU2\run\interface.py", line 279, in run_command
raise Exception , message
Exception: Path = D:\DESIGNS\DSN_001\DIRECT\,
Command = C:\SU2\SU2_CFD config_CFD.cfg
SU2 process returned error '1'
ERROR: Cannot find value AOA in given map.
Please check the name of the variable in the config file.
I am looking forward to your rely ,thank you

[btracey]

I'm not really sure to be honest. I've worked to make the config parsing better for the upcoming release (whatever the problem is, your error will be better soon).

My guess would be that your angle of attack variable is set to AoA. Could you try changing that to AOA and see if it fixes the issue?

[454514566@qq.com]

Originally Posted by btracey

I'm not really sure to be honest. I've worked to make the config parsing better for the upcoming release (whatever the problem is, your error will be better soon).

My guess would be that your angle of attack variable is set to AoA. Could you try changing that to AOA and see if it fixes the issue?

Hi ,Brendan Tracey,thank your reply fastly,I am sorry that I do not understand your answer fully,if I delete the last order( 102, 1.0 | airfoil ),it has not error,but the AOA is fixed.
But I want to optimize the airfoil when the CL is fixed,solving the AOA and airfoil shape that is similar to reverse design,because I see the AOA can set the design variable,so the AOA can be setted as incidence angle in 3D wing.

[economon]

Hi,

Just to follow up on this thread.. you might want to give the new fixed AoA mode in SU2 V3.2 a try for your problem. This doesn't directly address the shape design problem, but you may find it interesting. You can activate it with the following new options:

% Activate fixed lift mode (specify a CL instead of AoA, NO/YES)
FIXED_CL_MODE= NO
%
% Target coefficient of lift for fixed lift mode (0.0 by default)
TARGET_CL= 0.0
%
% Damping factor for fixed CL mode (0.1 by default)
DAMP_FIXED_CL= 0.1

All the best,
Tom

[454514566@qq.com]

Originally Posted by economon

Hi,

Just to follow up on this thread.. you might want to give the new fixed AoA mode in SU2 V3.2 a try for your problem. This doesn't directly address the shape design problem, but you may find it interesting. You can activate it with the following new options:

% Activate fixed lift mode (specify a CL instead of AoA, NO/YES)
FIXED_CL_MODE= NO
%
% Target coefficient of lift for fixed lift mode (0.0 by default)
TARGET_CL= 0.0
%
% Damping factor for fixed CL mode (0.1 by default)
DAMP_FIXED_CL= 0.1

All the best,
Tom

Sorry,I recompute the file lately,I hope that the airfoil shape and AOA is changing at the every time calculation,but I change the file form your advise,
it seem change the AOA only,the mesh is not change,and the situation is unstable.
My file is
inv_NASA0712.zip
If you need my mesh, call me and sent the mesh by email.
Thank you

[economon]

Hi,

Just to clear this up (my original post may have been a little misleading). The new fixed lift mode is not set up for use with optimization, rather, I just wanted to alert you of it's existence so that you can experiment with it for your analysis problem. Apologies for any confusion,

T

[454514566@qq.com]

Originally Posted by economon

Hi,

Just to clear this up (my original post may have been a little misleading). The new fixed lift mode is not set up for use with optimization, rather, I just wanted to alert you of it's existence so that you can experiment with it for your analysis problem. Apologies for any confusion,

T

So the AOA can be not as the DV for the SU2 in the shape optimization??

[hoyos98]

You can find the optimal airfoil shape (minimizing the drag) but fulfilling a target lift coefficient on this website: https://aeroptimal.com/airfoil . The methodology is under review in journals. The CFD method employs OpenFOAM.